EP4183066B1 - Verfahren zur datenübertragung durch ein raumfahrzeug mit einem laser-sendemodul - Google Patents

Verfahren zur datenübertragung durch ein raumfahrzeug mit einem laser-sendemodul Download PDF

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Publication number
EP4183066B1
EP4183066B1 EP22782915.7A EP22782915A EP4183066B1 EP 4183066 B1 EP4183066 B1 EP 4183066B1 EP 22782915 A EP22782915 A EP 22782915A EP 4183066 B1 EP4183066 B1 EP 4183066B1
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Prior art keywords
spacecraft
laser
calibration
sight
data
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English (en)
French (fr)
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EP4183066A1 (de
EP4183066C0 (de
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Mehdi Ghezal
Emmanuel GIRAUD
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Airbus Defence and Space SAS
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Airbus Defence and Space SAS
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    • H—ELECTRICITY
    • H04—ELECTRIC COMMUNICATION TECHNIQUE
    • H04B—TRANSMISSION
    • H04B10/00—Transmission systems employing electromagnetic waves other than radio-waves, e.g. infrared, visible or ultraviolet light, or employing corpuscular radiation, e.g. quantum communication
    • H04B10/11—Arrangements specific to free-space transmission, i.e. transmission through air or vacuum
    • H04B10/118—Arrangements specific to free-space transmission, i.e. transmission through air or vacuum specially adapted for satellite communication
    • H—ELECTRICITY
    • H04—ELECTRIC COMMUNICATION TECHNIQUE
    • H04B—TRANSMISSION
    • H04B10/00—Transmission systems employing electromagnetic waves other than radio-waves, e.g. infrared, visible or ultraviolet light, or employing corpuscular radiation, e.g. quantum communication
    • H04B10/50—Transmitters
    • H04B10/501—Structural aspects
    • H04B10/503—Laser transmitters
    • H—ELECTRICITY
    • H04—ELECTRIC COMMUNICATION TECHNIQUE
    • H04B—TRANSMISSION
    • H04B10/00—Transmission systems employing electromagnetic waves other than radio-waves, e.g. infrared, visible or ultraviolet light, or employing corpuscular radiation, e.g. quantum communication
    • H04B10/60—Receivers
    • H04B10/61—Coherent receivers
    • H04B10/616—Details of the electronic signal processing in coherent optical receivers
    • H04B10/6165—Estimation of the phase of the received optical signal, phase error estimation or phase error correction

Definitions

  • the present invention belongs to the field of data transmission by spacecraft, such as observation satellites, and more particularly relates to a method of transmitting data by a spacecraft in moving orbit, said data corresponding to acquired images by an observation instrument of the spacecraft, as well as a spacecraft for implementing such a transmission method.
  • Earth observation missions carried out by a spacecraft consist of acquiring images of parts of the Earth's surface, that is to say taking pictures of it. Such acquisitions are for example carried out in response to customer requests, and serve as a basis for the production of final composite images.
  • such a spacecraft follows a moving orbit around the Earth in order to carry out acquisitions during its flight over the Earth's surface.
  • it comprises an observation instrument associated with a predetermined spatial resolution as well as an optical line of sight.
  • an optical line of sight forms the outgoing part of the optical path of the observation instrument, and points towards the surface of the Earth during image acquisitions.
  • the acquisition method generally used for the observation of a terrestrial area is that known under the name of “strip scanning” (or “pushbroom” in the Anglo-Saxon literature).
  • a line sensor In such a “pushbroom” mode, a line sensor successively acquires a plurality of line images and the image of the complete terrestrial area, called the “composite image”, is obtained by combining said line images.
  • THE patent EP 3488540 B1 proposes an observation instrument operating in “pushbroom” mode, comprising a line sensor and a laser emission module both located at a focal plane of an optic of the observation instrument. Such arrangements are advantageous since they make it possible to reduce to a certain extent the quantity of equipment to be carried on board the spacecraft, since the same optics are used by both the line sensor and the laser emission module.
  • the present invention aims to remedy all or part of the drawbacks of the prior art, in particular those set out above, by proposing a solution which makes it possible to improve the pointing of the laser emission module by limiting the equipment to be on board. in the spacecraft.
  • the pointing error is essentially a two-dimensional unknown, for example modeled by an attitude error in roll (for example around a roll axis collinear with a speed vector of the spacecraft in inertial reference) and a pitch attitude error (e.g. around a pitch axis perpendicular to an orbit plane of the spacecraft), etc.
  • attitude error in roll for example around a roll axis collinear with a speed vector of the spacecraft in inertial reference
  • a pitch attitude error e.g. around a pitch axis perpendicular to an orbit plane of the spacecraft
  • the senor of the observation instrument is a matrix sensor, that is to say a sensor comprising a plurality of lines of acquisition cells and a plurality of columns of acquisition cells.
  • a matrix sensor that is to say a sensor comprising a plurality of lines of acquisition cells and a plurality of columns of acquisition cells.
  • the laser emission module and the matrix sensor use (at least in part) the same optics of the observation instrument.
  • the laser emission module is for example located in the focal plane of the optics like the matrix sensor, or in a secondary focal plane obtained by duplication of the focal plane so that the focal plane and the secondary focal plane overlap optically.
  • the laser emission module can also be located in an intermediate focal plane corresponding to part of the optics of the observation instrument. For example, if the observation instrument has Korsch type optics, the intermediate focal plane can correspond to the focal plane of the associated Cassegrain.
  • the pointing error is essentially the same for the matrix sensor and the laser emitting module. Consequently, the matrix sensor, used to acquire the images which are then emitted by the laser emission module, is advantageously also used to estimate the pointing error of the laser emission module.
  • the matrix sensor acquires a (2D) calibration image of a calibration zone for which reference data are available.
  • These reference data associated with the calibration zone correspond to data deducible from an image representing the calibration zone, and correspond to the values expected for these data in the absence of pointing error. Therefore, by comparing the reference data to the calibration image (i.e. to the corresponding data values as deduced from the calibration image), it is possible to observe the pointing error and estimate it.
  • the pointing error thus estimated is then used to correct the pointing of the laser line of sight.
  • the orientation of the laser sight line is then controlled taking into account the estimated pointing error, to point the laser sight line towards the considered set point.
  • the same matrix sensor which is used both to acquire the images emitted by the laser emission module (that is to say the images acquired as part of the mission observation of the spacecraft) and to acquire the calibration images used to calibrate the pointing error of the laser line of sight of said laser emission module. It is therefore not necessary to provide hardware means dedicated to the calibration of the pointing error, such as another matrix sensor dedicated to the acquisition of calibration images to estimate the pointing error.
  • the transmission method may also include, optionally, one or more of the following characteristics, taken in isolation or in all technically possible combinations.
  • the reference data comprise a theoretical image of the calibration zone.
  • the laser reception module is integrated into a ground station and the calibration zone is a terrestrial zone associated with said laser reception module.
  • the calibration zone is a sky zone so that the calibration image represents stars located in the field of view of the matrix sensor and the data of reference are determined from a catalog of stars.
  • the laser reception module is embarked in another spacecraft in Earth orbit.
  • the calibration zone comprises a light source of predetermined position and the reference data comprises a theoretical position of the light source in the calibration image, the pointing error being determined by comparison of 'an estimated position of the light source in the calibration image with the theoretical position.
  • the light source is collocated with the laser reception module.
  • the calibration zone being an area on the surface of the Earth, the spacecraft having a ground speed V ground and the observation instrument being associated with a spatial resolution R s following a direction of travel
  • the acquisition of the calibration image is carried out during a so-called immobilization duration greater than R s /V ground during which the attitude of the spacecraft is controlled so that an imprint on the ground of the field of vision is held stationary on the surface of the Earth.
  • the immobilization duration is significantly greater than R s /V ground (factor 100 or even 1000).
  • the steps of the method are iterated for the transmission of data to the same laser reception module, so as to alternate calibration phases and emission phases, the calibration phases carrying out the estimation of the pointing error and the transmission phases carrying out the transmission of data to the same reception module.
  • the pointing error can possibly vary during the same data transmission, for example due to a thermoelastic deformation of the structure of the spacecraft which varies over time.
  • the spacecraft may also optionally include one or more of the following characteristics, taken in isolation or in all technically possible combinations.
  • the capacity Ct is greater than 0.8 N ⁇ m and/or the capacity C r is greater than 0.8 N ⁇ m.
  • the attitude control means comprise at least one reaction wheel recovering electrical energy and/or at least one gyroscopic actuator.
  • the observation instrument comprises at least two fixed mirrors as a spacecraft reference, and the module laser emission emits laser radiation following the laser line of sight via at least two mirrors of the observation instrument.
  • the observation instrument comprises Korsch optics.
  • FIG 1 schematically represents a spacecraft in orbit moving (that is to say a non-geostationary orbit) around the Earth 80.
  • the spacecraft is placed in a circular orbit, with an altitude preferably less than 2000 km, or even less than 1000 km.
  • an altitude preferably less than 2000 km, or even less than 1000 km.
  • the spacecraft is a satellite 10.
  • spacecraft space shuttle, probe, etc.
  • Said satellite 10 moves in its circular orbit around the Earth 80 with a ground speed V ground .
  • the satellite 10 travels in a circular orbit of altitude substantially equal to 500 km with a ground speed substantially equal to 7 km.s -1 .
  • the satellite 10 includes an observation instrument 20 for acquiring images of parts of the surface of the Earth 80.
  • the observation instrument 20 comprises, in a manner known per se, acquisition means configured to capture the optical flow coming from the surface of the Earth 80.
  • the observation instrument 20 comprises at least one sensor matrix 24 comprising a plurality of acquisition cells (pixels) organized in several rows and several columns.
  • each line extends transversely to the direction of scrolling, while each column extends substantially in said direction of scrolling.
  • the observation instrument 20 also includes optics comprising one or more mirrors arranged to reflect the optical flow coming from the surface of the Earth 80 towards said matrix sensor 24, arranged at a focal plane PF of the optics of said observation instrument 20.
  • the observation instrument 20 is associated with an optical line of sight 21 (part a) of the figure 1 ).
  • Said optical line of sight 21 forms the outgoing part of the optical path of the observation instrument 20 and meets the surface of the Earth 80 at a point called “ground point” S.
  • the optical line of sight 21 is defined as corresponding to the optical path starting from the center of the matrix sensor 24.
  • the optical line of sight 21 is fixed as a reference satellite, that is to say that the orientation of the optical line of sight 21 relative to a body 11 of the satellite cannot be modified and follows the movement of the satellite 10 when the attitude of said satellite 10 is modified.
  • the observation instrument 20 is also associated with a fixed field of view in satellite reference.
  • This fixed field of vision corresponds to the angular aperture of the observation instrument 20 and the acquisition cells of the matrix sensor 24.
  • the field of vision of the observation instrument 20 forms a footprint of length L following the direction of travel.
  • the length of the footprint can vary with the incidence of the optical line of sight 21 on the surface of the Earth 80, and the length L corresponds to the minimum length of the footprint, which is obtained with a Nadir pointing of the optical line of sight 21.
  • image acquisitions are not necessarily carried out with Nadir pointing and can be carried out with any pointing.
  • the observation instrument 20 is associated with a predetermined spatial resolution R s .
  • the spatial resolution corresponds to the size, for example in meters, of the smallest object which can be detected in a scene represented by an image acquired by the observation instrument 20.
  • the spatial resolution R s is less than two meters (2 m), or even less than one meter (1 m).
  • the spatial resolution can vary with the incidence of the optical line of sight 21 on the surface of the Earth 80.
  • the spatial resolution R s corresponds here to the spatial resolution of the observation instrument 20 obtained with a Nadir pointing (and at the perigee of the orbit in the case of a elliptical orbit).
  • the spatial resolution R s is less than the length L and is preferably significantly less than said length L, for example by a factor of at least 5000, or even at least 10000.
  • the satellite 10 also includes a laser transmission module 30 for transmitting data in the form of laser radiation.
  • the data to be transmitted in the form of laser radiation notably comprises images of the surface of the Earth 80 acquired by the observation instrument 20. These data are transmitted to a laser reception module 40, which is found in a ground station on the surface of the Earth 80 in the non-limiting example of the figure 1 .
  • a laser reception module 40 can be provided, for example distributed on the surface of the Earth 80 to increase the opportunities for transferring data to the ground. It is also possible to consider one or more laser reception modules 40 on board other satellites, for example in geostationary orbit (GEO).
  • GEO geostationary orbit
  • the connection between the laser emission module 30 and a laser reception module 40 is generally referred to as a “laser connection”.
  • the laser emission module 30 comprises, in a manner known per se, a source of laser radiation and means adapted to modulate the laser radiation as a function of the data to be transmitted.
  • the laser emission module 30 is associated with a laser line of sight 31 (part b) of the figure 1 ) fixed in satellite reference, such as the optical sight line 21.
  • the optical 21 and laser 31 sight lines are linked in that the laser emission module 30 is integrated into the observation instrument 20, so that the laser emission module 30 uses all or part of an optic of the observation instrument 20.
  • the laser emission module 30 is located at the focal plane PF of the observation instrument 20 , like the matrix sensor 24.
  • Other examples are also described below. It should be noted that the optical sight lines 21 and 31 can be different from each other, or confused with each other, depending on the examples.
  • the optical 21 and laser 31 sight lines are essentially subject to the same pointing error.
  • FIG. 2 schematically represents the main stages of a method 50 of transmitting data by the satellite 10, by means of the laser emission module 30.
  • the transmission method 50 comprises a step S50 of acquisition, by the matrix sensor 24, of an image of a so-called calibration zone, called calibration image.
  • the calibration zone is a predetermined zone, for example an area on the surface of the Earth 80 or an area of sky against a background of stars, for which reference data can be obtained during a step S51.
  • reference data examples of which will be described in more detail below, are data deducible from an image representing the calibration zone, and correspond to the values expected for these data in the absence of pointing error.
  • a predetermined set point is associated with the calibration zone, and the attitude of the satellite 10 can be controlled to orient the optical line of sight 21 towards this set point.
  • the matrix sensor 24 then acquires the calibration image, which is therefore a representation of the calibration zone produced by pointing the optical line of sight 21 towards the set point.
  • the point actually targeted by the optical line of sight 21 is distinct from the set point, and the calibration image differs from what it would be in the absence of pointing error.
  • the acquisition of the calibration image is carried out during a so-called immobilization duration greater than R s /V sol during which the attitude of the satellite 10 is controlled so that a footprint on the ground of the field of view is kept stationary on the surface of the Earth 80.
  • the immobilization duration is significantly greater than R s /V sol (factor 100 or even 1000), because this contributes to significantly improving the SNR of the calibration image.
  • the emission method 50 comprises then a step S52 of determining the pointing error of the laser line of sight by comparison of the calibration image and the reference data.
  • this comparison can be direct (for example if the reference data correspond to a reference image of the calibration zone considered) or indirect (for example a pre-processing must be applied beforehand to the calibration image and/or reference data).
  • the reference data being representative of the expected values in the absence of pointing error, the difference between the reference data and the corresponding data deduced from the calibration image is therefore representative of the pointing error, and this deviation can be used to estimate said pointing error.
  • the pointing error can be estimated in different forms, for example as a roll attitude error and a pitch attitude error, or as a vector connecting the set point to the point actually aimed due to the pointing error, etc.
  • the transmission method 50 includes a step S53 of controlling the pointing of the satellite 10 as a function of the pointing error, to point the laser line of sight 31 towards a laser reception module 40.
  • the orientation of the laser line of sight 31 is controlled to compensate for the estimated pointing error.
  • the transmission method 50 then comprises a step S54 of transmitting data by the laser emission module 30 to said laser reception module 40, via the laser link.
  • the data may include, in particular, one or more images acquired by the matrix sensor 24 as part of the observation mission of the satellite 10.
  • the attitude of the satellite 10 is controlled in order to maintain the laser line of sight 31 oriented towards the laser reception module 40. If we consider for example a laser reception module 40 integrated into a ground station stationary on the surface of the Earth 80, this therefore means that the pointing of the laser line of sight 31 must be kept stationary on the surface of the Earth 80 for the entire duration of data transmission step S54.
  • keep stationary we mean that the satellite attitude setpoint is determined to maintain the point targeted by the laser line of sight 31 substantially stationary on the laser reception module 40, itself stationary on the surface of the Earth. 80, despite the movement of the satellite 10.
  • substantially immobile we mean that the objective of the attitude control is to maintain the point targeted by the laser line of sight 31 immobile, but that this can vary slightly during the duration of the transmission step S54, due for example to attitude control errors and/or measurement noise. Due to such attitude control, it is understood that the incidence of the laser line of sight 31 on the surface of the Earth 80 varies during the duration of the emission step S54. For example, if the incidence of the laser line of sight 31 is substantially normal to the surface of the Earth 80 at the start of the emission step S54, then the incidence of said laser line of sight 31 on said surface of the Earth 80 will be slightly oblique at the end of the transmission step S54, to compensate for the movement of the satellite 10. Thus, the attitude instruction given varies during the transmission step S54 and aims in particular to stop the scrolling of the laser sight line 31 on the ground.
  • the calibration zone is located near the laser reception module 40, in order to limit the duration of the maneuver necessary to go from pointing the optical line of sight 21 to the point of setpoint associated with the calibration zone at a pointing of the laser line of sight 31 on the laser reception module 40.
  • the setpoint associated with the calibration zone and the reception module laser 40 can be confused, in particular if the optical line of sight 21 and the laser line of sight 31 are confused.
  • the calibration zone is preferably a terrestrial zone associated with said ground station, that is to say a zone predetermined terrestrial comprising said ground station or near it. If the laser reception module 40 is on board another spacecraft, for example another satellite in GEO orbit, the calibration zone is preferably a sky zone associated with said other satellite, that is to say a zone of sky in the background of the other satellite relative to the observation instrument 20 or offset relative to said other satellite.
  • the satellite 10 has, in preferred embodiments, high torque forming capacities, in particular with respect to the inertia of said satellite 10, in order to be able to carry out possible maneuvers very quickly, and thus limit the duration between the estimation of the pointing error and the transmission of the data.
  • the drift of the pointing error may be non-negligible during the time necessary to transmit all of the data to be transmitted, for example if the pointing error drifts quickly and/or if the quantity of data to be transmitted is important. In this case, it may be necessary to track the pointing error.
  • the data are transmitted to the same laser reception module 40, so that the pointing control step S53 of each transmission phase aims to orient the laser line of sight 31 towards this laser reception module 40.
  • the calibration phases can all acquire calibration images of the same calibration zone, or acquire calibration images of different calibration zones.
  • FIG. 3 schematically represents the satellite 10 in the process of transmitting data to a laser reception module 40 on the surface of the Earth 80.
  • the transmission of data is broken down into two transmission phases PE 1 and PE 2 preceded by respective calibration phases PC 1 and PC 2 .
  • the transmission method 50 firstly comprises a first calibration phase PC 1 aimed at making a first estimate of the pointing error.
  • the matrix sensor 24 acquires a first calibration image of a first calibration zone, obtains the first associated reference data and deduces the first estimate of the pointing error.
  • the transmission method 50 then comprises a first transmission phase PE 1 aimed at transmitting data to the laser reception module 40.
  • the laser line of sight 31 is oriented towards the laser reception module 40 taking into account the first estimation of the pointing error, and data are transmitted by the laser emission module 30 while maintaining the pointing of the laser line of sight 31 stationary on the laser transmission module. laser reception 40.
  • the transmission method 50 then comprises a second calibration phase PC 2 aimed at carrying out a second estimate of the error of pointing.
  • the matrix sensor 24 acquires a second calibration image of a second calibration zone (which may be the same as the first calibration zone), obtains the second associated reference data ( if necessary) and deduces the second estimate of the pointing error.
  • the transmission method 50 then comprises a second transmission phase PE 2 aimed at continuing the transmission of the data to the laser reception module 40.
  • the laser line of sight 31 is oriented towards the laser reception module 40 taking into account the second estimation of the pointing error, and data is transmitted by the laser emission module 30 while maintaining the pointing of the laser line of sight 31 stationary on the laser reception module 40.
  • the reference data associated with a calibration zone correspond to a theoretical image of said calibration zone.
  • This theoretical image represents the image of the calibration zone which should have been obtained by the satellite 10, in the absence of pointing error, during the acquisition of the calibration image.
  • the registration of the calibration image with the theoretical image makes it possible to match the pixels of these images, that is to say makes it possible to identify the pixels of these images which represent the same portion of the calibration zone. Comparison of the positions in the images of pixels representing the same portion of the calibration zone can be used to estimate the pointing error affecting the acquisition of the calibration image.
  • the reference data includes a theoretical position in the calibration image of an element characteristic of the calibration zone.
  • characteristic element we mean an element which can be detected in the calibration image, for example because it must present a characteristic shape in the calibration image or, preferably, because it is reflected in the calibration image by one or more pixels presenting characteristic values.
  • the theoretical position in the calibration image of this characteristic element taking into account the position of the satellite 10 in its orbit and the set point targeted during this acquisition, and possibly a digital terrain model of the calibration zone.
  • the actual position of the characteristic element in the calibration image can be estimated by detecting said characteristic element in the calibration image.
  • the comparison of the theoretical position and the actual position of the characteristic element in the calibration image can then be used to estimate the pointing error on the acquisition of the calibration image.
  • the characteristic element is a light source, which emits light radiation which leads to characteristic values of the pixel(s) which represent said light source in the calibration image.
  • the light radiation emitted by the light source is for example radiation in the visible wavelength range.
  • the calibration zone is not illuminated by the Sun, high intensity light radiation, even in the visible wavelength range, can be detected. in the calibration image, especially since its theoretical position in said calibration image is known.
  • the light radiation emitted by the light source is preferably radiation in the range of lengths of non-visible waves, for example infrared radiation such as near infrared radiation (“near infrared” or NIR).
  • infrared radiation such as near infrared radiation (“near infrared” or NIR).
  • NIR radiation near infrared radiation
  • the matrix sensor 24 is also sensitive in the non-visible wavelength range used by the light source.
  • the light radiation emitted by the light source is laser radiation.
  • laser radiation can be detected more precisely in the calibration image, which makes it possible to improve the precision of the estimation of the pointing error.
  • the light source is a laser transmission module, which can be co-located with the laser reception module 40, used to transmit data to the satellite 10 (which optionally includes a laser reception module).
  • the calibration zone is a sky zone so that the calibration image represents stars located in the field of view of the matrix sensor 24, and the reference data are determined from a catalog of stars.
  • a star catalog includes information making it possible to know the positions of certain stars, for example the ephemeris of these stars.
  • the matrix sensor 24 of the observation instrument 20 is therefore used as a stellar sensor.
  • the pointing error can then be estimated, for example, by comparing the estimated attitude with the attitude measured by other attitude sensors of the satellite 10.
  • the pointing error can then be estimated, for example, by comparing the actual positions of said stars in the calibration image with their theoretical positions.
  • FIG. 4 schematically represents an exemplary embodiment of a satellite 10 for implementing the transmission method 50.
  • the satellite 10 comprises two solar generators 12, 13 arranged on respective opposite faces of a body 11 of said satellite 10.
  • the observation instrument 20 is arranged on a face connecting said faces carrying the solar generators 12, 13.
  • the satellite 10 also includes attitude control means (not shown in the figures), such as inertial actuators.
  • Said attitude control means have a pitching torque formation capacity Ct and a rolling torque formation capacity C r .
  • Ct pitching torque formation capacity
  • C r rolling torque formation capacity
  • the pitch inertia of the satellite 10 is designated It and the rolling inertia of the satellite 10 is designated I r .
  • the satellite 10 further comprises a processing circuit (not shown in the figures), which controls the operation of the observation instrument 20, the laser emission module 30 and the attitude control means.
  • the processing circuit comprises for example one or more processors and storage means (magnetic hard disk, electronic memory, optical disk, etc.) in which a computer program product is stored, in the form of a set of program code instructions to be executed to implement the different steps of the transmission method 50.
  • the processing circuit comprises one or more programmable logic circuits (FPGA, PLD, etc.), and/or one or more specialized integrated circuits (ASIC, etc.), and/or a set of discrete electronic components , etc., suitable for carrying out all or part of said steps of the emission method 50.
  • the processing circuit corresponds to means configured in software (specific computer program product) and/or hardware (FPGA, PLD, ASIC, discrete electronic components, etc.) to carry out all or part of the stages of the emission process 50, by a control adapted from the observation instrument 20, the laser emission module 30 and the attitude control means.
  • the satellite 10 can also include, in a conventional manner, other elements such as sensors (star sensor, gyrometer, etc.), which can also be connected to the processing circuit.
  • sensors star sensor, gyrometer, etc.
  • the ratio C t /I t is greater than 0.01 s -2 , or even greater than 0.018 s -2 .
  • the pitching torque formation capacity Ct is greater than 0.8 Newton meters (N ⁇ m) and the pitching inertia It is less than 80 kg ⁇ m 2 .
  • the pitching torque formation capacity Ct is greater than 1 N ⁇ m and the pitching inertia It is less than 60 kg ⁇ m 2 .
  • the attitude control means comprise one or more reaction wheels that recover electrical energy.
  • Such electrical energy-recovering reaction wheels are known from the demand for patent EP 2247505 A1 , notably.
  • reaction wheels that recover electrical energy is particularly advantageous for the following reasons.
  • first of all it should be noted that, to provide a high pitch (resp. roll) torque capacity (greater than 0.8 N ⁇ m or even greater than 1 N ⁇ m) by means of reaction wheels, it is necessary to use fairly massive reaction wheels, which tends to increase the pitching inertia It (resp. I r ) of the satellite 10.
  • reaction wheels recovering electrical energy in particular as described in the patent application EP 2247505 A1 , it is possible to reduce the required mass, for equivalent torque capacity, compared to reaction wheels that do not recover electrical energy.
  • the reaction wheels recover electrical energy, the electrical energy requirements of the satellite 10 are reduced.
  • the attitude control means comprise one or more gyroscopic actuators (“Control Moment Gyroscope” or CMG in the English literature).
  • gyroscopic actuators are particularly advantageous because they have a high ratio (torque / mass capacity). Thus, it is possible to have a high torque capacity without penalizing the inertia of the satellite 10.
  • the observation instrument 20 comprises at least one matrix sensor 24, for example of the CMOS type (acronym for the Anglo-Saxon expression “Complementary Metal-Oxide-Semiconductor”). It should however be noted that the observation instrument 20 may include several matrix sensors 24. For example, several matrix sensors can be used to acquire images in different respective wavelengths (red, green, blue, near infrared, etc.), etc.
  • the observation instrument 20 includes a Bayer filter.
  • a Bayer filter in a manner known per se, makes it possible to acquire color images in respective different wavelengths red, green and blue ("Red, Green, Blue” or RGB in the Anglo-Saxon literature) with the same matrix sensor 24. Such arrangements make it possible to simplify the observation instrument 20.
  • part a) of the Figure 5 represents a matrix sensor 24 comprising several sets of acquisition cells, sensitive respectively in the red wavelengths (designated by R on the figure 8 ), green (G), blue (B) and near infrared (NIR).
  • the matrix sensor 24 has 16x16 acquisition cells, and can be extended to a larger number of acquisition cells by repeating for example the pattern of 4 x 4 acquisition cells located in the upper left part (surrounded by a broken line).
  • Part b) of the Figure 5 represents the spectral responses of the different filters associated respectively with the red (R), green (G), blue (B) and near infrared (NIR) wavelengths. It is also possible, according to other examples, to have several focal planes, including a focal plane with a matrix sensor and a classic Bayer filter (forming a Bayer matrix) and another focal plane with at least one other sensor matrix and another filter, for example near infrared. Finally, nothing prevents breaking up the focal plane following respective filters adapted to a particular mission or using a single focal plane without a filter.
  • the observation instrument 20 is configured to successively activate acquisition cells during the acquisition of an image.
  • acquisition mode is known as “rolling shutter acquisition mode” in the Anglo-Saxon literature.
  • Such arrangements particularly suited to very large matrix sensors, make it possible to limit the quantity of data that must be processed simultaneously by the observation instrument 20, since the acquisition cells are not all activated simultaneously.
  • observation instrument 20 may include other elements, such as for example optics comprising one or more mirrors, one or more lenses, a support structure, electronic components, etc.
  • the observation instrument 20 comprises at least two fixed mirrors as a spacecraft reference, and the laser emission module 30 emits laser radiation following the laser line of sight 31 via at least two mirrors of the observation instrument 20.
  • the mirrors are used to successively reflect an optical flow received from the surface of the Earth 80 in the direction of the matrix sensor 24, all or part of these mirrors are also used to reflect in the opposite direction the laser radiation emitted by the laser emission module 30.
  • the observation instrument 20 and the laser emission module 30 are therefore structurally linked to each other , and are therefore subject to the same pointing errors. They can be seen as corresponding to one and the same equipment for pointing control operations.
  • FIGS. 6 to 8 represent non-limiting examples of embodiments in which mirrors of the observation instrument 20 are used by the laser emission module 30. It should be noted that the figures 6 to 8 are shown to scale, but only for the different mirrors and their respective positions.
  • the optics of the observation instrument 20 are of the Korsch type, which makes it possible to have both great compactness and low mass, with a high focal length.
  • Korsch's optics feature a 480mm M1 mirror, a 160mm M2 mirror, an M3 folding mirror and an M4 mirror. The incident optical flow is therefore successively reflected by the mirror M1, the mirror M2 (through the mirror M1), the folding mirror M3 and the mirror M4, until reaching the focal plane PF at which the matrix sensor is located. 24 of observation instrument 20.
  • the focal plane may include for example one or more matrix sensors, each matrix sensor comprising 14192 x 10140 acquisition cells (pixels), for example according to the IMX 411 model marketed by Sony® .
  • each matrix sensor comprising 14192 x 10140 acquisition cells (pixels), for example according to the IMX 411 model marketed by Sony® .
  • pixels 14192 x 10140 acquisition cells
  • body 11 of the satellite has a dimension of 1.53 x 1.14 x 1.0 m.
  • Two solar generators 12, 13 of 1 m 2 provide a power of 250 W sufficient for the needs of the satellite.
  • the east attitude of the satellite 10 is for example controlled around a reference attitude in which the X axis is collinear with the roll axis and the Y axis is collinear with the pitch axis, in which case the inertias I xx and I yy correspond respectively to the rolling inertia I r and the pitch inertia It.
  • the laser emission module 30 is located opposite the mirror M2, at the level of an intermediate focal plane associated with the mirrors M1 and M2.
  • the laser radiation emitted by the laser emission module 30 is therefore successively reflected by the mirror M2 then by the mirror M1 from which it is reflected along the laser line of sight 31.
  • the laser emission module 30 is located at the focal plane PF of the observation instrument 20, at a location on the focal plane PF which is offset relative to the matrix sensor 24.
  • the laser radiation emitted by the module d The laser emission 30 is therefore successively reflected by the mirrors M4, M3, M2 and M1 from which it is reflected along the laser line of sight 31.
  • the optical lines of sight 21 and 31 are necessarily different from one another. on the other, and the laser line of sight 31 is not in the field of vision of the matrix sensor 24.
  • the calibration image aims to detect a light source which corresponds to a module laser emission module on the ground co-located with the laser reception module 40, then it is necessary to carry out a maneuver between the acquisition of the calibration image (the laser emission module on the ground and the laser reception module 40 being located in the field of vision of the matrix sensor 24) and the transmission of data (the laser line of sight 31, outside the field of vision of the matrix sensor 24, being oriented towards the laser reception module 40).
  • a ratio C t /I t (resp. C r /I r ) greater than 0.01 s -2 , or even greater than 0.018 s -2 , makes it possible to reduce the duration of this maneuver.
  • the laser emission module 30 is located at a secondary focal plane PS, which corresponds to a duplication of the focal plane PF of the observation instrument 20.
  • a focal plane duplication element ED is arranged on the path of the radiation between the focal plane PF and the mirror M4.
  • the duplicating element ED may include a mirror or dichroic plate configured to reflect one among the laser radiation of the laser emission module 30 and the optical flow coming from the observed scene and to transmit without reflection the other among the laser radiation and the optical flow.
  • it is the laser radiation which is reflected, so that the laser radiation is successively reflected by the duplication element ED then the mirrors M4, M3, M2 and M1.
  • optical sight lines 21 and 31 can be confused. If we consider for example the case where the calibration image aims to detect a light source which corresponds to a laser emission module on the ground co-located with the laser reception module 40, then the fact of having optical sight lines 21 and laser 31 combined makes it possible to greatly reduce the requirements in terms of maneuver between the acquisition of the calibration image and the transmission of the data to the laser reception module 40.
  • the optics of the observation instrument 20 may in particular include a number of mirrors different from the number of mirrors (4) represented on the figures 6 to 8 .
  • the optics of the observation instrument may include two mirrors, for example arranged like the mirrors M1 and M2 of the Figure 6 , the laser emission module 30 and the matrix sensor 24 being for example both located at the focal plane of the observation instrument 20 (which corresponds in this case to the intermediate focal plane of the Figure 6 ).

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  • Engineering & Computer Science (AREA)
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  • Signal Processing (AREA)
  • Astronomy & Astrophysics (AREA)
  • General Physics & Mathematics (AREA)
  • Optics & Photonics (AREA)
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  • Optical Radar Systems And Details Thereof (AREA)

Claims (15)

  1. Verfahren (50) zum Senden von Daten aus einem Raumfahrzeug (10), das die Erde (80) auf einer Umlaufbahn umrundet, wobei das Raumfahrzeug aufweist:
    - ein Beobachtungsinstrument (20), das eine Optik und einen in einer Fokalebene (PF) der Optik angeordneten Matrixsensor (24) aufweist, um Bilder im Rahmen einer Beobachtungsmission des Raumfahrzeugs zu erfassen, wobei das Beobachtungsinstrument ein zugehöriges Sichtfeld hat, das bezogen auf das Raumfahrzeug ortsfest ist und von dem Matrixsensor (24) definiert ist,
    - ein Laseremissionsmodul (30) mit einer zugehörigen Laservisierlinie (31), die bezogen auf das Raumfahrzeug ortsfest ist, wobei das Laseremissionsmodul in der Fokalebene oder in einer sekundären Fokalebene (PS) der Optik oder in einer Zwischenfokalebene eines Teils der Optik angeordnet ist,
    wobei das Verfahren folgende Schritte aufweist:
    - (S50) Erfassen eines Bilds einer Kalibrierungszone, Kalibrierungsbild genannt, mithilfe des Matrixsensors,
    - (S51) Abrufen von Referenzdaten, die mit der Kalibrierungszone verknüpft sind,
    - (S52) Bestimmen eines Ausrichtungsfehlers der Laservisierlinie durch Vergleichen des Kalibrierungsbilds und der Referenzdaten,
    - (S53) Steuerung der Ausrichtung des Raumfahrzeugs durch Korrigieren des Ausrichtungsfehlers, um die Laservisierlinie hin zu einem Laserempfangsmodul auszurichten,
    - (S54) Senden von Daten aus dem Laseremissionsmodul mit Ziel Laserempfangsmodul, wobei die Daten ein oder mehrere von dem Matrixsensor (24) im Rahmen der Beobachtungsmission des Raumfahrzeugs erfasste Bilder aufweisen.
  2. Verfahren (50) nach Anspruch 1, wobei die Referenzdaten ein theoretisches Bild der Kalibrierungszone aufweisen.
  3. Verfahren (50) nach Anspruch 2, wobei das Laserempfangsmodul in einer Bodenstation eingebaut ist und die Kalibrierungszone eine terrestrische Zone ist.
  4. Verfahren (50) nach Anspruch 1 oder 2, wobei die Kalibrierungszone eine Himmelszone ist, so dass das Kalibrierungsbild Sterne darstellt, die sich in dem Sichtfeld des Matrixsensors befinden, und die Referenzdaten aus einem Sternenkatalog bestimmt werden.
  5. Verfahren (50) nach Anspruch 4, wobei das Laserempfangsmodul an Bord eines anderen Raumfahrzeugs ist, das sich in einer Erdumlaufbahn befindet.
  6. Verfahren (50) nach Anspruch 1, wobei die Kalibrierungszone eine Lichtquelle in einer vorgegebenen Position aufweist und die Referenzdaten eine theoretische Position der Lichtquelle in dem Kalibrierungsbild aufweisen, wobei der Ausrichtungsfehler durch Vergleichen einer geschätzten Position der Lichtquelle in dem Kalibrierungsbild mit der theoretischen Position bestimmt wird.
  7. Verfahren (50) nach Anspruch 6, wobei die Lichtquelle gemeinsam mit dem Laserempfangsmodul lokalisiert wird.
  8. Verfahren (50) nach einem der vorstehenden Ansprüche, wobei das Raumfahrzeug eine Umlaufgeschwindigkeit am Boden Vsol hat und das Beobachtungsinstrument eine zugehörige räumliche Auflösung Rs in einer Umlaufrichtung hat, wobei die Erfassung des Kalibrierungsbilds während einer sogenannten Unbeweglichkeitszeitdauer von mehr als Rs/Vsol durchgeführt wird, während der die Lage des Raumfahrzeugs derart gesteuert wird, dass ein Bodenabdruck des Sichtfelds auf der Erdoberfläche unbeweglich gehalten wird.
  9. Verfahren (50) nach einem der vorstehenden Ansprüche, wobei die Schritte des Verfahrens für das Senden von Daten hin zu einem gleichen Laserempfangsmodul derart iteriert werden, dass Kalibrierungsphasen (PC1, PC2) und Sendephasen (PE1, PE2) einander abwechseln, wobei die Kalibrierungsphasen die Schätzung des Ausrichtungsfehlers durchführen und die Sendephasen das Senden von Daten hin zu dem gleichen Laserempfangsmodul durchführen.
  10. Raumfahrzeug (10), das vorgesehen ist, in eine Umlaufbahn gebracht zu werden, auf der es die Erde (80) in einer Umlaufrichtung umrundet, aufweisend:
    - Einrichtungen zur Lagesteuerung des Raumfahrzeugs,
    - ein Beobachtungsinstrument (20) mit einem zugehörigen Sichtfeld, das bezogen auf das Raumfahrzeug ortsfest ist und von einem in einer Fokalebene (PF) einer Optik des Beobachtungsinstruments angeordneten Matrixsensor (24) definiert ist,
    - ein Laseremissionsmodul (30) mit einer zugehörigen Laservisierlinie (31), die bezogen auf das Raumfahrzeug ortsfest ist, wobei das Laseremissionsmodul in der Fokalebene oder in einer sekundären Fokalebene (PS) der Optik oder in einer Zwischenfokalebene eines Teils der Optik angeordnet ist,
    - Einrichtungen, die konfiguriert sind, ein Verfahren zum Senden von Daten nach einem der vorstehenden Ansprüche durchzuführen.
  11. Raumfahrzeug (10) nach Anspruch 10, wobei:
    - das Raumfahrzeug (10) eine Nickträgheit It hat und die Einrichtungen zur Lagesteuerung eine Nickmomentbildungskapazität Ct haben, wobei das Verhältnis Ct/It größer als 0,01 s-2 ist, und/oder
    - das Raumfahrzeug (10) eine Rollträgheit Ir hat und die Einrichtungen zur Lagesteuerung eine Rollmomentbildungskapazität Cr haben, wobei das Verhältnis Cr/Ir größer als 0,01 s-2 ist.
  12. Raumfahrzeug (10) nach Anspruch 11, wobei die Kapazität Ct größer als 0,8 N*m ist und/oder die Kapazität Cr größer als 0,8 N*m ist.
  13. Raumfahrzeug (10) nach einem der Ansprüche 10 bis 12, wobei die Einrichtungen zur Lagesteuerung mindestens ein Schwungrad zur Rückgewinnung von elektrischer Energie und/oder mindestens einen gyroskopischen Aktuator aufweisen.
  14. Raumfahrzeug (10) nach einem der Ansprüche 10 bis 13, wobei das Beobachtungsinstrument mindestens zwei Spiegel (M1, M2, M3, M4) aufweist, die bezogen auf das Raumfahrzeug ortsfest sind, und das Laseremissionsmodul über die mindestens zwei Spiegel des Beobachtungsinstruments einen Laserstrahl entlang der Laservisierlinie emittiert.
  15. Raumfahrzeug (10) nach Anspruch 14, wobei das Beobachtungsinstrument eine Korsch-Optik aufweist.
EP22782915.7A 2021-09-08 2022-09-05 Verfahren zur datenübertragung durch ein raumfahrzeug mit einem laser-sendemodul Active EP4183066B1 (de)

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PCT/FR2022/051669 WO2023037067A1 (fr) 2021-09-08 2022-09-05 Procédé d'émission de données par un engin spatial comportant un module d'émission laser

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Publication number Priority date Publication date Assignee Title
EP2247505B1 (de) * 2008-02-11 2011-10-26 Astrium Sas Betätigungsvorrichtung zur veränderung der fluglage eines raumfahrzeuges

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FR2960313B1 (fr) * 2010-05-18 2012-07-27 Astrium Sas Procede de prise d'image
FR3058012B1 (fr) 2016-10-21 2020-01-10 Airbus Defence And Space Sas Systeme combine d'imagerie et de communication par signaux laser
CN110999129B (zh) * 2017-06-14 2023-08-15 穿梭科技私人投资有限公司 用于高速通信的系统和方法
CN111901032B (zh) * 2020-08-25 2021-09-07 中国科学院微小卫星创新研究院 一种一体化星载光学传感器系统

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Publication number Priority date Publication date Assignee Title
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